US20060045741A1 - Methods and apparatus for cooling gas turbine engine rotor assemblies - Google Patents
Methods and apparatus for cooling gas turbine engine rotor assemblies Download PDFInfo
- Publication number
- US20060045741A1 US20060045741A1 US10/932,492 US93249204A US2006045741A1 US 20060045741 A1 US20060045741 A1 US 20060045741A1 US 93249204 A US93249204 A US 93249204A US 2006045741 A1 US2006045741 A1 US 2006045741A1
- Authority
- US
- United States
- Prior art keywords
- shank
- rotor
- rotor blade
- platform
- accordance
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
- 238000000034 method Methods 0.000 title claims abstract description 20
- 230000000712 assembly Effects 0.000 title description 4
- 238000000429 assembly Methods 0.000 title description 4
- 239000000112 cooling gas Substances 0.000 title description 2
- 238000001816 cooling Methods 0.000 claims abstract description 25
- 230000008878 coupling Effects 0.000 claims abstract description 18
- 238000010168 coupling process Methods 0.000 claims abstract description 18
- 238000005859 coupling reaction Methods 0.000 claims abstract description 18
- 241000879887 Cyrtopleura costata Species 0.000 claims description 18
- 238000011144 upstream manufacturing Methods 0.000 claims description 16
- 238000010926 purge Methods 0.000 claims description 9
- 238000004513 sizing Methods 0.000 claims 1
- 239000007789 gas Substances 0.000 description 12
- 238000005336 cracking Methods 0.000 description 2
- 230000003647 oxidation Effects 0.000 description 2
- 238000007254 oxidation reaction Methods 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 230000008646 thermal stress Effects 0.000 description 2
- 239000000567 combustion gas Substances 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
- F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for cooling gas turbine engine rotor assemblies.
- At least some known rotor assemblies include at least one row of circumferentially-spaced rotor blades. Each rotor blade includes an airfoil that includes a pressure side, and a suction side connected together at leading and trailing edges. Each airfoil extends radially outward from a rotor blade platform to a tip, and also includes a dovetail that extends radially inward from a shank extending between the platform and the dovetail. The dovetail is used to couple the rotor blade within the rotor assembly to a rotor disk or spool. At least some known blades are hollow such that an internal cooling cavity is defined at least partially by the airfoil, platform, shank, and dovetail.
- During operation, because the airfoil portions of the blades are exposed to higher temperatures than the dovetail portions, temperature gradients may develop at the interface between the airfoil and the platform, and/or between the shank and the platform. Over time, thermal strain caused by such temperature gradients may induce compressive thermal stresses to the blade platform. Moreover, over time, the increased operating temperature of the platform may cause platform oxidation, platform cracking, and/or platform creep deflection, which may shorten the useful life of the rotor blade.
- To facilitate reducing the effects of the high temperatures in the platform region, at least some known rows of rotor blades are coupled to a rotor disk such that a predetermined gap is defined between adjacent blade platforms. The gap enables leakage of cooling air to circulate near the platform region. However, within known rotor blades, such gaps may provide only limited cooling to the rotor blade platforms.
- In one embodiment, a method for assembling a rotor assembly for gas turbine engine is provided. The method includes providing a first rotor blade that includes an airfoil, a platform, a shank, and a dovetail, wherein the airfoil extends radially outward from the platform, the shank extends radially inward from the platform, and the dovetail extends from the shank, forming a recess within a portion of the shank, coupling the first rotor blade to a rotor shaft using the dovetail, and coupling a second rotor blade to the rotor shaft such that a shank cavity is defined between the first and second rotor blade shanks, such that, during operation, cooling air may enter and pressurize the shank cavity through the recessed portion.
- In another embodiment, a rotor blade for a gas turbine engine is provided. The rotor blade includes a platform, an airfoil extending radially outward from the platform, a shank extending radially inward from the platform, and a dovetail extending radially inward from the shank, wherein at least a portion of the shank is recessed to facilitate increasing pressure of cooling air supplied to a shank cavity defined adjacent said shank during engine operation.
- In a further embodiment, a gas turbine engine rotor assembly is provided. The rotor assembly includes a rotor shaft, and a plurality of circumferentially-spaced rotor blades coupled to the rotor shaft wherein each rotor blade includes an airfoil extending radially outward from a platform, a shank extending radially inward from the platform, and a dovetail extending from the shank for coupling the rotor blade to the rotor shaft, each shank includes a pair of opposing sidewalls that extend axially between an upstream sidewall and a downstream sidewall, the plurality of rotor blades are circumferentially-spaced such that a shank cavity is defined between each pair of adjacent rotor blades, at least a portion of the rotor blade shank upstream sidewall is recessed such that the shank cavity may be pressurized during engine operation.
-
FIG. 1 is schematic illustration of a gas turbine engine; -
FIG. 2 is an enlarged perspective view of a rotor blade that may be used with the gas turbine engine shown inFIG. 1 ; -
FIG. 3 is an enlarged perspective view of the rotor blade shown inFIG. 2 and viewed from the underside of the rotor blade; and -
FIG. 4 is a front view illustrating a relative orientation of the circumferential spacing between the rotor blade shown inFIG. 2 and other rotor blades when coupled within the gas turbine engine shown inFIG. 1 . -
FIG. 1 is a schematic illustration of an exemplarygas turbine engine 10 coupled to anelectric generator 16. In the exemplary embodiment,gas turbine system 10 includes acompressor 12, aturbine 14, andgenerator 16 arranged in a single monolithic rotor orshaft 18. In an alternative embodiment,shaft 18 is segmented into a plurality of shaft segments, wherein each shaft segment is coupled to an adjacent shaft segment to formshaft 18.Compressor 12 supplies compressed air to acombustor 20 wherein the air is mixed withfuel 22 supplied thereto. In one embodiment,engine 10 is a 7FA+e gas turbine engine commercially available from General Electric Company, Greenville, S.C. - In operation, air flows through
compressor 12 and compressed air is supplied tocombustor 20.Combustion gases 28 fromcombustor 20propels turbines 14.Turbine 14 rotatesshaft 18,compressor 12, andelectric generator 16 about alongitudinal axis 30. -
FIG. 2 is an enlarged perspective view of arotor blade 40 that may be used with gas turbine engine 10 (shown inFIG. 1 ) viewed from afirst side 42 ofrotor blade 40.FIG. 3 is an enlarged perspective view ofrotor blade 40 and viewed from an underside ofrotor blade 10.FIG. 4 is a front view and illustrates a relative orientation of circumferential spacing defined between circumferentiallyadjacent rotor blades 40, whenblades 40 are coupled within a rotor assembly, such as turbine 14 (shown inFIG. 1 ). In the exemplary embodiment,blade 40 has been modified to include the features described herein. More specifically, whenrotor blades 40 are coupled within the rotor assembly, apredetermined platform gap 48 is defined between the circumferentiallyadjacent rotor blades 40. - When coupled within the rotor assembly, each
rotor blade 40 is coupled to a rotor disk (not shown) that is rotatably coupled to a rotor shaft, such as shaft 18 (shown inFIG. 1 ). In an alternative embodiment,blades 40 are mounted within a rotor spool (not shown). In the exemplary embodiment, circumferentiallyadjacent blades 40 are identical and each extends radially outward from the rotor disk and includes anairfoil 60, aplatform 62, ashank 64, and adovetail 66. In the exemplary embodiment,airfoil 60,platform 62,shank 64, anddovetail 66 are collectively known as a bucket. - Each
airfoil 60 includesfirst sidewall 70 and asecond sidewall 72.First sidewall 70 is convex and defines a suction side ofairfoil 60, andsecond sidewall 72 is concave and defines a pressure side ofairfoil 60.Sidewalls edge 74 and at an axially-spacedtrailing edge 76 ofairfoil 60. More specifically, airfoiltrailing edge 76 is spaced chord-wise and downstream fromairfoil leading edge 74. - First and
second sidewalls blade root 78 positionedadjacent platform 62, to anairfoil tip 80.Airfoil tip 80 defines a radially outer boundary of aninternal cooling chamber 84 is defined withinblades 40. More specifically,internal cooling chamber 84 is bounded withinairfoil 60 betweensidewalls platform 62 and throughshank 64 and intodovetail 66. -
Platform 62 extends betweenairfoil 60 andshank 64 such that eachairfoil 60 extends radially outward from eachrespective platform 62. Shank 64 extends radially inwardly fromplatform 62 to dovetail 66, anddovetail 66 extends radially inwardly fromshank 64 to facilitate securingrotor blades Platform 62 also includes an upstream side orskirt 90 and a downstream side orskirt 92 that are connected together with a pressure-side edge 94 and an opposite suction-side edge 96. Whenrotor blades 40 are coupled within the rotor assembly,platform gap 48 is defined between adjacentrotor blade platforms 62, and accordingly is known as a platform gap. -
Shank 64 includes a substantiallyconcave sidewall 120 and a substantially convex sidewall (not shown) connected together at anupstream sidewall 124 and a downstream sidewall (not shown) ofshank 64. Accordingly,shank sidewall 120 is recessed with respect to upstream anddownstream sidewalls 124 and 126, respectively, such that whenblades 40 are coupled within the rotor assembly, ashank cavity 128 is defined between adjacentrotor blade shanks 64. - In the exemplary embodiment, a
forward angel wing 130 and anaft angel wing 132 each extend outwardly from respectiveshank sides 124 and 126 to facilitate sealing forward and aft angel wing buffer cavities (not shown) defined within the rotor assembly. In addition, a forwardlower angel wing 134 also extends outwardly fromshank side 124 to facilitate sealing betweenblades 40 and the rotor disk. More specifically, forwardlower angel wing 134 extends outwardly fromshank 64 betweendovetail 66 andforward angel wing 130. - To facilitate increasing a pressure within
shank cavity 128, in the exemplary embodiment,shank sidewall 124 may be modified to include a recessed or scallopedportion 160 formed radially inward from forwardlower angel wing 134. Recessedportion 160 is sized and oriented to permit a predetermined amount of cooling airflow into shank cavity. In the exemplary embodiment,recessed portion 160 is substantially parallel tolongitudinal axis 30. Accordingly, whenadjacent rotor blades 40 are coupled within the rotor assembly, recessedportion 160 enables additional cooling air flow intoshank cavity 128 to facilitate increasing an operating pressure withinshank cavity 128. As such,recessed portion 160 facilitates maintaining a sufficient back flow margin for platform cooling usingplatform gap 48. - Generally, during engine operation,
bucket pressure side 42 generally operates at higher temperatures than rotorblade suction side 44. Cooling air enteringshank cavity 128 through shank sidewall recessedportion 160 facilitates maintaining a sufficient back flow margin withinshank cavity 128 such that at least a portion of the cooling air withinshank 128 may be channeled through platform undercutpurge slot 170 and throughplatform gap 48. As the cooling air is forced outward throughslot 170 andplatform gap 48,platform 62 is convectively cooled to facilitate reducing the operating temperature ofplatform 62 such that thermal strains induced toplatform 62 are also reduced. - The above-described rotor blades provide a cost-effective and reliable method for supplying cooling air to facilitate reducing an operating temperature of the rotor blade platform. More specifically, through cooling flow, thermal stresses induced within the platform, and the operating temperature of the platform is facilitated to be reduced. Accordingly, platform oxidation, platform cracking, and platform creep deflection is also facilitated to be reduced. As a result, the rotor blade cooling circuit facilitates extending a useful life of the rotor assembly and improving the operating efficiency of the gas turbine engine in a cost-effective and reliable manner.
- Exemplary embodiments of rotor blades and rotor assemblies are described above in detail. The rotor blades are not limited to the specific embodiments described herein, but rather, components of each rotor blade may be utilized independently and separately from other components described herein. For example, each rotor blade cooling circuit component can also be used in combination with other rotor blades, and is not limited to practice with
only rotor blade 40 as described herein. Rather, the present invention can be implemented and utilized in connection with many other blade and cooling circuit configurations. - While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.
Claims (20)
Priority Applications (4)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US10/932,492 US7189063B2 (en) | 2004-09-02 | 2004-09-02 | Methods and apparatus for cooling gas turbine engine rotor assemblies |
DE102005040304A DE102005040304A1 (en) | 2004-09-02 | 2005-08-24 | Method and apparatus for cooling gas turbine engine rotors |
JP2005248565A JP2006070899A (en) | 2004-09-02 | 2005-08-30 | Method and device for cooling gas turbine engine rotor assembly |
CNA2005100990892A CN1743646A (en) | 2004-09-02 | 2005-09-02 | Methods and apparatus for cooling gas turbine engine rotor assemblies |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US10/932,492 US7189063B2 (en) | 2004-09-02 | 2004-09-02 | Methods and apparatus for cooling gas turbine engine rotor assemblies |
Publications (2)
Publication Number | Publication Date |
---|---|
US20060045741A1 true US20060045741A1 (en) | 2006-03-02 |
US7189063B2 US7189063B2 (en) | 2007-03-13 |
Family
ID=35943393
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US10/932,492 Active 2024-12-04 US7189063B2 (en) | 2004-09-02 | 2004-09-02 | Methods and apparatus for cooling gas turbine engine rotor assemblies |
Country Status (4)
Country | Link |
---|---|
US (1) | US7189063B2 (en) |
JP (1) | JP2006070899A (en) |
CN (1) | CN1743646A (en) |
DE (1) | DE102005040304A1 (en) |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090324424A1 (en) * | 2007-09-28 | 2009-12-31 | Daniel Tragesser | Air cooled bucket for a turbine |
US20100003127A1 (en) * | 2007-09-28 | 2010-01-07 | Ian Reeves | Air cooled bucket for a turbine |
EP4166757A1 (en) * | 2021-10-15 | 2023-04-19 | Rolls-Royce plc | Bladed disc |
EP4166755A1 (en) * | 2021-10-15 | 2023-04-19 | Rolls-Royce plc | Bladed disc |
Families Citing this family (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8038405B2 (en) * | 2008-07-08 | 2011-10-18 | General Electric Company | Spring seal for turbine dovetail |
US8226365B2 (en) * | 2009-04-22 | 2012-07-24 | General Electric Company | Systems, methods, and apparatus for thermally isolating a turbine rotor wheel |
US8540486B2 (en) * | 2010-03-22 | 2013-09-24 | General Electric Company | Apparatus for cooling a bucket assembly |
GB2486488A (en) | 2010-12-17 | 2012-06-20 | Ge Aviat Systems Ltd | Testing a transient voltage protection device |
US8951014B2 (en) | 2011-03-15 | 2015-02-10 | United Technologies Corporation | Turbine blade with mate face cooling air flow |
US8876479B2 (en) | 2011-03-15 | 2014-11-04 | United Technologies Corporation | Damper pin |
US8967973B2 (en) * | 2011-10-26 | 2015-03-03 | General Electric Company | Turbine bucket platform shaping for gas temperature control and related method |
US8827643B2 (en) * | 2011-10-26 | 2014-09-09 | General Electric Company | Turbine bucket platform leading edge scalloping for performance and secondary flow and related method |
US8893507B2 (en) | 2011-11-04 | 2014-11-25 | General Electric Company | Method for controlling gas turbine rotor temperature during periods of extended downtime |
US9039382B2 (en) | 2011-11-29 | 2015-05-26 | General Electric Company | Blade skirt |
US10247015B2 (en) * | 2017-01-13 | 2019-04-02 | Rolls-Royce Corporation | Cooled blisk with dual wall blades for gas turbine engine |
US10934865B2 (en) * | 2017-01-13 | 2021-03-02 | Rolls-Royce Corporation | Cooled single walled blisk for gas turbine engine |
US11840940B2 (en) | 2021-03-09 | 2023-12-12 | Mechanical Dynamics And Analysis Llc | Turbine blade tip cooling hole supply plenum |
Citations (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2603453A (en) * | 1946-09-11 | 1952-07-15 | Curtiss Wright Corp | Cooling means for turbines |
US2915279A (en) * | 1953-07-06 | 1959-12-01 | Napier & Son Ltd | Cooling of turbine blades |
US4726735A (en) * | 1985-12-23 | 1988-02-23 | United Technologies Corporation | Film cooling slot with metered flow |
US5020970A (en) * | 1989-07-13 | 1991-06-04 | Dresser-Rand Company | Fluid-handling, bladed rotor |
US5281097A (en) * | 1992-11-20 | 1994-01-25 | General Electric Company | Thermal control damper for turbine rotors |
US6273683B1 (en) * | 1999-02-05 | 2001-08-14 | Siemens Westinghouse Power Corporation | Turbine blade platform seal |
US6428270B1 (en) * | 2000-09-15 | 2002-08-06 | General Electric Company | Stage 3 bucket shank bypass holes and related method |
US6431833B2 (en) * | 1999-09-24 | 2002-08-13 | General Electric Company | Gas turbine bucket with impingement cooled platform |
US6478540B2 (en) * | 2000-12-19 | 2002-11-12 | General Electric Company | Bucket platform cooling scheme and related method |
US20050095128A1 (en) * | 2003-10-31 | 2005-05-05 | Benjamin Edward D. | Methods and apparatus for cooling gas turbine engine rotor assemblies |
Family Cites Families (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPS6463605A (en) * | 1987-09-04 | 1989-03-09 | Hitachi Ltd | Gas turbine moving blade |
JPH02140403A (en) * | 1988-11-18 | 1990-05-30 | Toshiba Corp | Mounting support structure for turbine rotor blade |
JP3502133B2 (en) * | 1993-12-28 | 2004-03-02 | 株式会社日立製作所 | Gas turbine and its rotor blade |
JP3040656B2 (en) * | 1994-05-12 | 2000-05-15 | 三菱重工業株式会社 | Gas Turbine Blade Platform Cooling System |
JP3040660B2 (en) * | 1994-06-06 | 2000-05-15 | 三菱重工業株式会社 | Gas Turbine Blade Platform Cooling Mechanism |
-
2004
- 2004-09-02 US US10/932,492 patent/US7189063B2/en active Active
-
2005
- 2005-08-24 DE DE102005040304A patent/DE102005040304A1/en not_active Ceased
- 2005-08-30 JP JP2005248565A patent/JP2006070899A/en active Pending
- 2005-09-02 CN CNA2005100990892A patent/CN1743646A/en active Pending
Patent Citations (10)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2603453A (en) * | 1946-09-11 | 1952-07-15 | Curtiss Wright Corp | Cooling means for turbines |
US2915279A (en) * | 1953-07-06 | 1959-12-01 | Napier & Son Ltd | Cooling of turbine blades |
US4726735A (en) * | 1985-12-23 | 1988-02-23 | United Technologies Corporation | Film cooling slot with metered flow |
US5020970A (en) * | 1989-07-13 | 1991-06-04 | Dresser-Rand Company | Fluid-handling, bladed rotor |
US5281097A (en) * | 1992-11-20 | 1994-01-25 | General Electric Company | Thermal control damper for turbine rotors |
US6273683B1 (en) * | 1999-02-05 | 2001-08-14 | Siemens Westinghouse Power Corporation | Turbine blade platform seal |
US6431833B2 (en) * | 1999-09-24 | 2002-08-13 | General Electric Company | Gas turbine bucket with impingement cooled platform |
US6428270B1 (en) * | 2000-09-15 | 2002-08-06 | General Electric Company | Stage 3 bucket shank bypass holes and related method |
US6478540B2 (en) * | 2000-12-19 | 2002-11-12 | General Electric Company | Bucket platform cooling scheme and related method |
US20050095128A1 (en) * | 2003-10-31 | 2005-05-05 | Benjamin Edward D. | Methods and apparatus for cooling gas turbine engine rotor assemblies |
Cited By (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090324424A1 (en) * | 2007-09-28 | 2009-12-31 | Daniel Tragesser | Air cooled bucket for a turbine |
US20100003127A1 (en) * | 2007-09-28 | 2010-01-07 | Ian Reeves | Air cooled bucket for a turbine |
US8052395B2 (en) | 2007-09-28 | 2011-11-08 | General Electric Company | Air cooled bucket for a turbine |
US8147188B2 (en) | 2007-09-28 | 2012-04-03 | General Electric Company | Air cooled bucket for a turbine |
EP4166757A1 (en) * | 2021-10-15 | 2023-04-19 | Rolls-Royce plc | Bladed disc |
EP4166755A1 (en) * | 2021-10-15 | 2023-04-19 | Rolls-Royce plc | Bladed disc |
US11814980B2 (en) | 2021-10-15 | 2023-11-14 | Rolls-Royce Plc | Bladed disc |
Also Published As
Publication number | Publication date |
---|---|
DE102005040304A1 (en) | 2006-05-24 |
JP2006070899A (en) | 2006-03-16 |
US7189063B2 (en) | 2007-03-13 |
CN1743646A (en) | 2006-03-08 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US7600972B2 (en) | Methods and apparatus for cooling gas turbine engine rotor assemblies | |
US7147440B2 (en) | Methods and apparatus for cooling gas turbine engine rotor assemblies | |
US6984112B2 (en) | Methods and apparatus for cooling gas turbine rotor blades | |
US7189063B2 (en) | Methods and apparatus for cooling gas turbine engine rotor assemblies | |
US6923616B2 (en) | Methods and apparatus for cooling gas turbine engine rotor assemblies | |
US7878763B2 (en) | Turbine rotor blade assembly and method of assembling the same | |
US7090466B2 (en) | Methods and apparatus for assembling gas turbine engine rotor assemblies | |
US8562286B2 (en) | Dead ended bulbed rib geometry for a gas turbine engine | |
US6921246B2 (en) | Methods and apparatus for assembling gas turbine nozzles | |
US20120045337A1 (en) | Turbine bucket assembly and methods for assembling same | |
CN102400717B (en) | Turbine blade platform cooling systems | |
US7458779B2 (en) | Gas turbine or compressor blade | |
US10655485B2 (en) | Stress-relieving pocket in turbine nozzle with airfoil rib | |
US7413409B2 (en) | Turbine airfoil with weight reduction plenum | |
US7597542B2 (en) | Methods and apparatus for controlling contact within stator assemblies | |
US7296966B2 (en) | Methods and apparatus for assembling gas turbine engines | |
KR20080001638A (en) | High performance turbine buckets and engines and turbines including same |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: GENERAL ELECTRIC COMPANY, NEW YORK Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:HONKOMP, MARK STEVEN;REEL/FRAME:015770/0888 Effective date: 20040901 |
|
FEPP | Fee payment procedure |
Free format text: PAYOR NUMBER ASSIGNED (ORIGINAL EVENT CODE: ASPN); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY |
|
STCF | Information on status: patent grant |
Free format text: PATENTED CASE |
|
FPAY | Fee payment |
Year of fee payment: 4 |
|
FPAY | Fee payment |
Year of fee payment: 8 |
|
MAFP | Maintenance fee payment |
Free format text: PAYMENT OF MAINTENANCE FEE, 12TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1553); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY Year of fee payment: 12 |
|
AS | Assignment |
Owner name: GE INFRASTRUCTURE TECHNOLOGY LLC, SOUTH CAROLINA Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:GENERAL ELECTRIC COMPANY;REEL/FRAME:065727/0001 Effective date: 20231110 |